EP1739309A1 - Mehrstufiger Turbomaschinenkompressor - Google Patents

Mehrstufiger Turbomaschinenkompressor Download PDF

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Publication number
EP1739309A1
EP1739309A1 EP06291053A EP06291053A EP1739309A1 EP 1739309 A1 EP1739309 A1 EP 1739309A1 EP 06291053 A EP06291053 A EP 06291053A EP 06291053 A EP06291053 A EP 06291053A EP 1739309 A1 EP1739309 A1 EP 1739309A1
Authority
EP
European Patent Office
Prior art keywords
ferrule
wall
annular
blades
casing
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP06291053A
Other languages
English (en)
French (fr)
Other versions
EP1739309B1 (de
Inventor
Gilles Alain Marie Charier
Franck Christian Conan
Thomas Julien Roland Earith
Marc Leandri
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA SAS
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Filing date
Publication date
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Application filed by SNECMA SAS filed Critical SNECMA SAS
Publication of EP1739309A1 publication Critical patent/EP1739309A1/de
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel

Definitions

  • the invention relates to a multi-stage turbomachine compressor, in particular a high-pressure compressor, and to an airplane turbojet or turboprop engine equipped with this compressor.
  • a high-pressure turbojet or turboprop compressor comprises a plurality of compression stages each comprising an annular row of rotor blades rotating inside a stationary casing and mounted on a shaft of the turbomachine, and a annular row of fixed blades for rectifying the flow of air, which are carried by the casing at their radially outer ends.
  • a double wall casing comprising a fixed outer wall and an inner wall which can move radially and which is connected to the outer wall by flexible suspension means or deformable, called "pins".
  • the compression stages are secured to one another, at least in groups of two, via the inner wall of the casing, which limits the possibilities of adjusting the radial clearances because these games are adjusted in the same way for compression stages that are solid, while the variations of these games differ from one stage to another.
  • the present invention is intended in particular to provide a simple, effective and economical solution to this problem.
  • It relates to a multi-stage turbomachine compressor, in particular a high-pressure compressor, in which it is possible to adjust the radial clearances of the different stages or at least some of the different stages independently for each stage.
  • a multi-stage turbomachine compressor comprising annular rows of moving blades rotating inside a double-walled casing and annular rows of fixed vanes mounted on an internal wall of the casing.
  • the inner wall of the housing comprises a plurality of rings substantially aligned end to end and suspended independently of one another to an outer wall of the housing, some of the rings each surrounding an annular row of blades and other ferrules each carrying an annular row of fixed vanes for recovery.
  • the different stages of compression are therefore, totally or at least with regard to certain of them, separated from each other, which allows to adjust the radial clearances of these floors independently for each floor, taking into account differences in mass, radial dimension and operating temperature between floors. This results in an improvement of the efficiency of the compressor and the operability of the turbomachine.
  • each annular row of blades is surrounded by a ferrule suspended from the outer wall of the casing independently of the other ferrules which surround the other annular rows of blades.
  • the ferrules are suspended from the outer wall of the housing by flexible or deformable means.
  • a first ferrule surrounding an annular row of blades is adjacent to a second ferrule carrying an annular row of fixed blades for straightening, one axial end of this second ferrule being connected to the first ferrule. by means ensuring a seal with respect to the flow of gas passing through the compressor.
  • the second axial end of the second ferrule is connected to a third ferrule surrounding a another annular row of blades by means providing a seal against the flow of gas passing through the compressor.
  • the second axial end of the second ferrule may in this case be connected to the outer wall of the housing by the suspension means of the third ferrule.
  • the second axial end of the second ferrule is spaced axially from a third ferrule surrounding another annular row of blades and is connected by its own suspension means to the outer wall of the housing.
  • an annular clearance axially separating the second ferrule and the third ferrule constitutes an air inlet passage in a cavity which is formed in the housing between the suspension means of the second ferrule and those of the third ferrule. and in which open air intake means carried by the outer wall of the housing, these air intake means being connected to other equipment of the turbomachine.
  • the invention allows, in a general manner, that the radial vibration modes of the rings surrounding the annular rows of moving blades of the compressor are set independently of one ring to the other, that is to say one floor to the other, for an optimization of the radial clearances between these ferrules and the annular rows of moving blades which turn inside these ferrules.
  • the invention also relates to an aircraft turbojet or turboprop engine, characterized in that it comprises a high-pressure compressor as described above.
  • the compressor 10 of FIG. 1 which illustrates the prior art, comprises a certain number of compression stages of which only three have been shown, each stage comprising an annular row of moving blades 12, the radially inner ends of which are fixed on a disc carried by a shaft of the turbomachine, and an annular row of fixed vanes 14 of rectification, arranged downstream of the annular row of moving blades 12 and whose radially outer ends are carried by a radially inner wall 16 of a double-walled cylindrical housing 18.
  • the inner cylindrical wall 16 of the casing 18 is suspended from the outer cylindrical wall 20 of this casing by flexible or deformable means 22, commonly known as "pins" in the art, the shapes, the masses and the rigidities of which are known to be adjusted the inner wall 16 of the casing best follows the radial vibrations of the rotor comprising the annular rows of moving blades 12.
  • the inner cylindrical wall 16 of the housing is formed of coaxial ferrules 24, which are aligned end to end and rigidly connected to each other via their annular flanges 26, which extend radially towards the outside and which are fixed to each other by appropriate means such as bolts.
  • the outer wall 20 of the housing 18 may be composed of ferrules arranged end to end and fixedly connected to each other by their outer annular flanges 28 by means of bolts or the like.
  • these radial clearances can be regulated independently of one compression stage to another, thanks to the fact that the rings which constitute the inner wall of the housing and which surround the annular rows of moving blades, are suspended independently of each other at the outer wall of the housing, for all compressor compressor stages or for at least a major part of them .
  • the inner wall of the housing 18 is formed of a succession of ferrules 30, 32 respectively, which are suspended from the outer wall 20 of the housing independently of each other by pins 34, 36 respectively, each ferrule 30 surrounding an annular row of blades 12 and each ferrule 32 carrying an annular row of fixed vanes 14.
  • the ferrule 30 is connected to the ferrule 32 situated downstream by means 38 providing an annular seal between these two ferrules with respect to the flow of gas passing through the compressor, which makes it possible to ensure the continuity of this flow of gas in the compressor and prevents the entry of air into the space between the inner and outer walls of the housing 18.
  • the suspensions independent of the various rings 30, 32 of the inner wall of the casing make it possible to adjust independently of one another the radial clearances J1 and J2 between the radially outer ends of the blades 12 and the rings 30 in each compression stage. .
  • the order of magnitude of these radial clearances is the tenth of a millimeter, the number of compression stages of a high-pressure compressor can be between 5 and 10 approximately, depending on the engines.
  • the ferrule 30 of the compression stage E2 which surrounds the annular row of moving blades 12 of this stage, is connected at its downstream end to a shell 32 which carries the annular row of fixed vanes 14 for straightening this stage. and whose downstream end is connected by suspension means 36 to the outer wall of the casing 18.
  • the ferrule 32 of the compression stage E2 is separated from the ferrule 30 of the next compression stage E3 by an axial annular clearance 42 which forms a gas passage between the inside of the compressor and a cavity 44 defined in the crankcase. 18 between the inner and outer walls of the latter on the one hand and between the means 36 for suspending the ferrule 32 of the previous stage E2 and the means 34 for suspending the ferrule 30 of the stage E3, somewhere else.
  • One or more air intakes 46 are formed in the outer wall of the casing 18 and open into this cavity 44, to supply air to the equipment of the turbomachine.
  • the downstream end of the ferrule 30 of the compression stage E3 is sealingly connected to the upstream end of a ferrule 32 carrying the fixed vanes 14 of this compression stage.
  • the downstream end of this ferrule 32 is connected sealingly, for example by interlocking, to the upstream end of the shell 30 of the next compression stage E4 which is connected by its suspension means 34 to the outer wall of the casing 18.
  • the ferrule 32 of the compression stage E3 is thus carried by the ferrule 30 of this compression stage and by the ferrule 30 of the next compression stage E4.
  • Another air intake 46 may be formed in the outer wall of the casing 18 and opens into a cavity 48 formed between the inner and outer walls of the housing 18 downstream of the suspension means 34 of the ferrule 30 of the stage E4.
  • the radial clearances of the compression stage E2 can be regulated independently of the radial clearances of the compression stage E1 and the following compression stages E3 and E4 while the games
  • the radial blades of the compression stages E3 and E4 are set independently, the blades 12 of these two stages having the same radial dimensions, the rings 30 of the stages E3 and E4 being secured to one another by the ferrule 32. from floor E3.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP06291053.4A 2005-06-29 2006-06-26 Mehrstufiger Turbomaschinenkompressor Revoked EP1739309B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0506610A FR2887939B1 (fr) 2005-06-29 2005-06-29 Compresseur multi-etages de turbomachine

Publications (2)

Publication Number Publication Date
EP1739309A1 true EP1739309A1 (de) 2007-01-03
EP1739309B1 EP1739309B1 (de) 2017-01-11

Family

ID=36282762

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06291053.4A Revoked EP1739309B1 (de) 2005-06-29 2006-06-26 Mehrstufiger Turbomaschinenkompressor

Country Status (4)

Country Link
US (1) US7651317B2 (de)
EP (1) EP1739309B1 (de)
FR (1) FR2887939B1 (de)
RU (1) RU2375607C2 (de)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2913051B1 (fr) 2007-02-28 2011-06-10 Snecma Etage de turbine dans une turbomachine
US11466593B2 (en) 2020-01-07 2022-10-11 Raytheon Technologies Corporation Double walled stator housing

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4522559A (en) * 1982-02-19 1985-06-11 General Electric Company Compressor casing
FR2688539A1 (fr) * 1992-03-11 1993-09-17 Snecma Stator de turbomachine comprenant des dispositifs de reglage de jeu entre le stator et les aubes du rotor.
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
US5630702A (en) * 1994-11-26 1997-05-20 Asea Brown Boveri Ag Arrangement for influencing the radial clearance of the blading in axial-flow compressors including hollow spaces filled with insulating material
US6390771B1 (en) * 1999-06-10 2002-05-21 Snecma Moteurs High-pressure compressor stator

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB856599A (en) 1958-06-16 1960-12-21 Gen Motors Corp Improvements relating to axial-flow compressors
US3314648A (en) 1961-12-19 1967-04-18 Gen Electric Stator vane assembly
US3511577A (en) * 1968-04-10 1970-05-12 Caterpillar Tractor Co Turbine nozzle construction
GB1501916A (en) 1975-06-20 1978-02-22 Rolls Royce Matching thermal expansions of components of turbo-machines
US6935836B2 (en) * 2002-06-05 2005-08-30 Allison Advanced Development Company Compressor casing with passive tip clearance control and endwall ovalization control
DE102004016222A1 (de) 2004-03-26 2005-10-06 Rolls-Royce Deutschland Ltd & Co Kg Anordnung zur selbsttätigen Laufspalteinstellung bei einer zwei- oder mehrstufigen Turbine

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4522559A (en) * 1982-02-19 1985-06-11 General Electric Company Compressor casing
FR2688539A1 (fr) * 1992-03-11 1993-09-17 Snecma Stator de turbomachine comprenant des dispositifs de reglage de jeu entre le stator et les aubes du rotor.
US5351478A (en) * 1992-05-29 1994-10-04 General Electric Company Compressor casing assembly
US5630702A (en) * 1994-11-26 1997-05-20 Asea Brown Boveri Ag Arrangement for influencing the radial clearance of the blading in axial-flow compressors including hollow spaces filled with insulating material
US6390771B1 (en) * 1999-06-10 2002-05-21 Snecma Moteurs High-pressure compressor stator

Also Published As

Publication number Publication date
FR2887939A1 (fr) 2007-01-05
US20090304498A1 (en) 2009-12-10
EP1739309B1 (de) 2017-01-11
FR2887939B1 (fr) 2016-09-30
US7651317B2 (en) 2010-01-26
RU2375607C2 (ru) 2009-12-10
RU2006123033A (ru) 2008-01-10

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